In the compressor portion of an aircraft gas turbine engine, atmospheric air is compressed to 10–50 times atmospheric pressure, and adiabatically heated to about 800°–1450° F. (425°–790° C.) in the process. This heated and compressed air is directed into a combustor, where it is mixed with fuel. The fuel is ignited, and the combustion process heats the gases to very high temperatures, in excess of 3000° F. (1650° C.). These hot gases pass through the turbine, where rotating turbine wheels extract energy to drive the fan and compressor of the engine, and the exhaust system, where the gases supply thrust to propel the aircraft, passing along the centerbody. To improve the efficiency of operation of the aircraft engine, combustion temperatures have been raised. Of course, as the combustion temperature is raised, steps must be taken to prevent thermal degradation of the materials forming the flow path for these hot gases of combustion.
Every aircraft gas turbine engine has a so-called High Pressure Turbine (HPT) to drive its compressor. The HPT sits just behind the compressor in the engine layout and experiences the highest temperature and pressure levels (nominally 2400° F. and 300 psia respectively) developed in the engine. The HPT also operates at very high speeds (10,000 RPM for large turbofans, 50,000 for small helicopter engines). In order to meet life requirements at these levels of temperature and pressure, HPT's today are always air-cooled and constructed from advanced alloys.
While a straight turbojet engine will usually have only one turbine section (an HPT), most engines today are of the turbofan, either high bypass turbofan or low bypass turbofan, or turboprop type and require one (and sometimes two) additional turbine(s) sections to drive a fan or a gearbox. The additional turbines are called the Low Pressure Turbines (LPT) and immediately follow the HPT in the engine layout. Since substantial pressure drop occurs across the HPT, the LPT operates with a much less energetic fluid and will usually require several stages (usually up to six) to extract the power required to perform its function.
In high-bypass fan turbofan engines the primary exhaust nozzle assembly consists of a centerbody and an outer barrel fairing. The turbine nozzle directs the primary gas aft and regulates the gas stream flow. The centerbody is aerodynamically shaped and has an outer surface that is located within the flow path of the hot core gas that has passed through the LPT. The outer surfaces of the outer fixed cowl provide a continuation of the fan flowpath. The nozzle outer barrel assembly is comprised of an outer cowl bolted to the inner barrel position called the primary nozzle. There are ports on the inner barrel to introduce engine low pressure recoup into the primary airstream. In low-bypass fan turbofan engines, the centerbody is not part of the primary exhaust nozzle assembly, but rather is positioned directly aft of the LPT and fore of the augmenter.
One well-known solution that has been undertaken to protect the metals that form the flow path for the hot gases of combustion, including those of the HPT, LPT, and primary nozzle assembly have included application of protective layers having low thermal conductivity. These materials are applied as thermal barrier coating systems (TBCs), typically comprising an overlying insulative ceramic top coat, typically a stabilized zirconia, and a bond coat that improves adhesion of the top coat to the substrate. These systems are known to improve the thermal performance of the underlying metals that form the flow path in the hot section of the engine. However, as temperatures of combustion have increased, even these TBCs have been found to be insufficient. While some modifications of the traditional flow path surfaces have been applied in the past, such as the application of materials over the TBC, these modifications have been directed to reducing the emissions of pollutants such as unburned hydrocarbons (UHC) and carbon monoxide (CO). One such modification is set forth in U.S. Pat. No. 5,355,668 to Weil, et al., assigned to the assignee of the present invention, which teaches the application of a catalyst such as platinum, nickel oxide, chromium oxide or cobalt oxide directly over the flow path surface of the thermal barrier coating of a turbine engine component. The catalyst layer is applied to selected portions of flow path surfaces to catalyze combustion of fuel. The catalytic material is chosen to reduce air pollutants such as unburned hydrocarbons (UHC) and carbon monoxide (CO) resulting from the combustion process. The catalytic layer is applied to a thickness of 0.001 to 0.010 inches and is somewhat rough and porous, having a surface roughness of about 100 to 250 micro inches, in order to enhance the surface area available to maximize contact with the hot gases in order to promote the catalytic reaction. The rough surface assists in creating some turbulence that promotes contact with the catalytic surface.
As temperatures of gas turbine engines have continued to increase, the combustion temperatures have become sufficiently high that even the best superalloy materials exhibit shortened lives due to thermal degradation. This is true even of the superalloys used for splash plates in high efficiency, advanced cycle turbine engines, which are prone to failure by thermal degradation. As combustion temperatures have increased, the impingement cooling and thermal barrier coatings have been inadequate to provide sufficient cooling to maintain component life without thermal degradation. Various attempts have been made to improve the resistance to thermal degradation, which have provided incremental improvements. These have included high temperature reflectors referred to as “Spray and Bake” coatings. These reflectors include platinum paints and platinum layers applied by chemical vapor deposition deposited over silicon dioxide (SiO2). These reflectors act by reflecting heat away from the splash plate rather than having the heat absorbed by the splash plate, conducted through the splash plate and then removed from the back side (or upstream side) of the splash plate by convection. Ideally, the heat is reflected back into the flow of combustion gases moving downstream into the turbine portion of the engine. However, these “Spray and Bake” coatings become ineffective as the temperatures rise above about 1600° F. (870° C.).
The prior art solutions are either directed to problems that are unrelated to the problem of high temperature degradation experienced by gas turbine components, such as the Weil patent, or provide different solutions to the problem of high temperatures resulting from the combustion process. What is needed is a cost effective coating system that can act as a reflector to assist in cooling a thin splash plate by reflecting radiative heat back into the combustion gas stream. The coating must be sufficiently thin so as not to increase the weight of the component substantially, yet reduce the radiative heat absorbed by the engine components so that the components experience less thermal degradation, which ultimately requires component replacement.
The present invention therefore provides a different approach to solving the problem of high temperatures experienced by gas turbine engine component.